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固定翼飞行器是一种能够在空中飞行的飞行器，其空气动力学性能对于飞行器的设计和性能具有重要影响。紧凑涡旋格方法是一种用于计算固定翼飞机空气动力学的数值方法，它通过将飞机的几何形状离散化为小的元素，并利用格点上的涡旋来模拟飞机表面的气流情况，从而计算飞机的升力、阻力和其他空气动力学性能。
使用紧凑涡旋格方法进行固定翼飞机空气动力学的研究可以帮助工程师和设计师更好地理解飞机的飞行特性，优化飞机的设计，提高飞机的性能和安全性。同时，这种方法也可以用于模拟飞机在不同飞行状态下的空气动力学性能，为飞行器的控制和飞行特性研究提供重要的数据支持。
QuadAir是一种用于计算飞机空气动力学性能的软件，可以用于模拟飞机在不同飞行状态下的气动力学力和力矩。使用QuadAir可以对Cessna 152进行空气动力学性能的计算和分析，包括升力、阻力、侧向力和俯仰力矩等参数的计算。
要使用QuadAir计算Cessna 152的空气动力学力和力矩，首先需要建立飞机的几何模型和飞行状态，然后进行数值模拟计算。通过这些计算，可以得到Cessna 152在不同飞行状态下的气动力学性能数据，为飞机的设计和性能评估提供重要的参考和支持。使用QuadAir进行Cessna 152的空气动力学力和力矩计算可以帮助工程师和设计师更好地理解飞机的气动特性，为飞机的设计和性能优化提供重要的数据支持。
综上，紧凑涡旋格方法是一种重要的研究工具，对于固定翼飞机空气动力学的研究具有重要意义，可以为飞机的设计、优化和性能提高提供重要的支持和指导。
部分代码：
%% Aircraft geometry parameters
Aircraft.r_CG_bo = [-0.375, 0, 0]; % position of CG with respect to apex of first wing, in Standard Frame (X: front, Y: right side, Z: down)
Aircraft.symmetry = [1]; % Indicates which wings have symetry about the X-Z plane
Aircraft.n_prt_wng = [3]; % Number of partitions per wing
Aircraft.flapped = [0 1 0]; % Indicates which partitions are flapped
Aircraft.flp_frac = [0 0.2 0] ; % Indicates the chord fraction occupied by the flap at eah partition. If partition is UNFLAPPED then set to 0.
Aircraft.spn = [6 7 4]*1.5/17; % Span (including the symmetric part if exists) of each partition of each wing
Aircraft.root_chrd = [2.2]*1/17; % Root chord of each wing
Aircraft.tpr_rto = [2.2 1.9 0.5]*1/2.2; % Taper ratio of each partition of each wing
Aircraft.swp_angle = [0 0 25]*pi/180; % Sweep angle of each partition of each wing
Aircraft.dih_angle = [0 0 10]*pi/180; % Dihedral angle of each partition of each wing
Aircraft.xyz_000 = [0 0 0]; % XYZ Position of the apex of each the wing, in Geometric Frame (X: back, Y: right side, Z: up). First one should be (0, 0, 0), to make it the reference point
Aircraft.twst_ang = [+2 +2 +2 +2]*pi/180; % Angle of incidence of each station of each wing (NOTE: each wing has #partitions+1 stations)
Aircraft.airfoil = {0 0 0}; % Airfoils are defined as an X-Y column matrix going continuously from TE to LE and all the way back to TE.
% Define airfoild drag polar with parabolic approximation: cd = cd_0 + cd_1*cl + cd_2*cl^2
% Assume NACA 2412 for main wing, and NACA 0009 for horizontal and vertical tail
Aircraft.cd_0 = [0.0151 0.0151 0.0151];
Aircraft.cd_1 = [-0.0126 -0.0126 -0.0126];
Aircraft.cd_2 = [0.0083 0.0083 0.0083];
Aircraft.wng_con_surf = [0]; % Indicating which wings are full control surfaces. The whole wing is rotated.
Aircraft.wng_con_surf_axis_rot = [0 1 0]; % Specifying the axis of rotation of each full wing control surface.
Aircraft.con_surf_group = [1 % Indicates grouping of control surfaces and symmetric/anti-symmetric relation
1];
%% Geometric Discretization Parameters
% The structure "geo_disc" holds the relevant geometric disretization parameters.
% UNFLAPPED part
geo_disc.spn_div(1,:,1) = [5 5 3]; % Number of span-wise divisions for the UN-FLAPPED part of each partition of each wing
geo_disc.chrd_div(1,:,1) = [5 4 3]; % Number of chordwise-wise divisions for the UN-FLAPPED part of each partition of each wing
% FLAPPED part
geo_disc.spn_div(1,:,2) = geo_disc.spn_div(:,:,1); % Set to be equal in FLAPPED and UN-FLAPPED parts.
geo_disc.chrd_div(1,:,2) = [0 3 0]; % Number of chordwise-wise divisions for the FLAPPED part of each partition of each wing
文章中一些内容引自网络，会注明出处或引用为参考文献，难免有未尽之处，如有不妥，请随时联系删除。
[1]王明振,李新颖,左仔滨,等.固定翼飞机水上迫降漂浮特性计算方法研究[J].航空科学技术, 2015(4):7.DOI:10.3969/j.issn.1007-5453.2015.04.015.
[2]刘烽.面向固定翼飞行器的大攻角飞行抗扰控制方法研究[J].[2024-01-15].
